Cooling techniques for high temperature engines and other components



Feb. 21, 1967 A. R. PARILLA COOLING TECHNIQUES FOR HIGH TEMPERATUREENGINES AND OTHER COMPONENTS 8 Sheets-Sheet 1 Filed April 12, 1963INVENTOR. ARTHUR R. PARILLA W050, l/ nljdw, film/n ATTORN EYS Feb. 21,1967 A. R. PARILLA COOLING TECHNIQUES FOR HIGH TEMPERATURE ENGINES ANDOTHER COMPONENTS 8 Sheets-Sheet 2 Filed April 12 1963 INVENTOR. ARTHURR. PARI LLA F f ,4 P MW wm m Feb. 21, 1967 PARlLLA 3,305,178

COOLING TECHNIQUES FOR HIGH TEMPERATURE ENGINES AND OTHER COMPONENTSFiled April 12, 1963 8 Sheets-Sheet 5 INVENTOR. ARTHU R R. PARILLAATTORNEYS Feb. 21, 1967 A. R. PARILLA 3,305,178

COOLING TECHNIQUES FOR HIGH TEMPERATURE ENGINES AND OTHER COMPONENTSFiled April 12, 1963 8 Sheets-Sheet 4 TA #N MP R M Mo A Wu Feb. 21, 1967A. R. PARILLA 3,305,178

COOLING TECHNIQUES FOR HIGH TEMPERATURE ENGINES AND OTHER COMPONENTSFiled April 12. 1963 s Sheets-Sheet 5 INVENTOR. ARTHUR R. PAR ILLAATTORN EYS A. R. PARILLA Feb. 21, 1967 COOLING TECHNIQUES FOR HIGHTEMPERATURE ENGINES AND OTHER COMPONENTS 8 Sheets-Sheet 6 Filed April12, 1963 29.23% \ESQQQ 453i HEQQQQA 3Q 3% m 3.8% k

INVENTOR. ARTHUR R. PARI LLA Mil ATTORNEYS Feb. 21, 1967 A. R. PARILLA3,305,178

COOLING TECHNIQUES FOR HIGH TEMPERATURE ENGINES AND OTHER COMPONENTS 8Sheets-Sheet 7 Filed April 12, 1963 INVENTOR. ARTHUR R- PARILLA /%ryan,flit/"44m! ATTORNEYS Feb. 21, 1967 A. R. PARILLA 3,305,178

COOLING TECHNIQUES FOR HIGH TEMPERATURE ENGINES AND OTHER COMPONENTSFiled April 12, 1963 8 Sheets-Sheet 8 1 1 I1 11 II 11 II 11 11 1/ 111/11 FIG.9

INVENTOR. ARTHUR R. PARI LLA BY 1 Mbrym, FM ML, ark/0% )VI/ILZ ATTORNEYSUnited States Patent 3,305,178 COOLING TECHNIQUES FOR HIGH TEMPERA- TUREENGINES AND OTHER COMPONENTS Arthur R. Parilla, 34 Crestview Road,Mountain Lakes, NJ. 07046 Filed Apr. 12, 1963, Ser. No. 273,265 19Claims. (Cl. 239-132.5)

This invention to cooling techniques for high temperature engines,particularly rocket engines. The invention is especially adapted to theeffective cooling of high performance rocket engines having combustionproduct temperatures exceeding by a substantial margin the meltingpoints of even the most rugged refractory metals.

While the invention will be described in terms of cooling techniques andnozzle structures for solid and liquid propellant rockets, it should beunderstood that several of the arrangements and processes according tothe invention will have applicability to the cooling of other hightemperature components as well.

Technological advances in rocket engine propellants have greatlyaggravated the problem of heat exchange in the exhaust section of theengine and particularly in the nozzle portion where heat transfer fromthe exhaust to the missile structure is maximum. The combustion productsof new propellants have temperatures in the range of 8000 F. and hencetheir use, with the attendant increase in rocket performance, dependsvitally on an effective coolant system.

When it is recalled that the best practical refractory metals havemelting points several thousand degrees below the exhaust temperaturescharacterizing these new fuels, it may be readily appreciated that thecooling problem is a complex and vexing one. If refractory metals are tobe used then the coolant or heat exchange system must keep wall membersformed of these materials at temperatures thousands of degrees belowexhaust temperatures. This means that a marked temperature differentialwill exist, causing massive heat transfer from exhaust to structure andrequiring an extraordinary coolant system.

It will also be appreciated upon reflection that the coolant system fora rocket must satisfy certain uniquely severe requirements. Since thrustis in the nature of a precious commodity, weight is obviously at apremium. A system which requires massive structures or more than a smallamount of coolant is clearly unacceptable. The system must operate atall expected flight attitudes and over severe ranges of accelerationmagnitudes and directions including gravity-free flight. For someapplications the system must be capable of discontinuous operation,terminating during in-flight shut down and commencing again withrestart. The system must also operate reliably in the presence of highamplitude shock and substantial ambient pressure variations ranging fromatmospheric to the vacuum of space. Notwithstanding these requirements,the system must have maximum reliability 1 and must therefore beextremely simple; the lives of operating personnel and the mammothinvestment in time and money depend on it.

It is accordingly a principal object of the invention to provideimproved heat exchange or cooling techniques for high performance rocketengines and other components operating at high temperatures.

It is a more specific object of the invention to provide techniques foreffectively cooling rocket engines having exhaust temperatures wellbeyond the melting points of known refractory metals, to the end thathigh performance fuels which generate such temperatures may effectivelyand reliably be employed.

A still further object of the invention is to provide cooling techniquescharacterized by the ability to func- "ice tion anew in space after atemporary shut down of engine operation and to provide coolingtechniques responsive to engine throttling.

Other objects of the invention are to provide improvements in the heatdissipating abilities of heat exchangers while concurrently effectingstructural simplifications thereof such as elimination of double wallstructures; to provide a cooling system which contributes to thrust; toprovide a system particularly adaptable to larger size rocket engines;to eliminate certain dependencies on ablative techniques; to eliminatethe need for external energy sources for preheating and coolant flowcontrol; to eliminate the need for regenerative cooling in liquidpropellant rockets together with the attendant pump system complexes;and to improve the control over and predictability of performance ofrocket engine coolant systems.

These and other objects and advantages of the invention such as thoserelating to nozzle fabrication processes Will be set forth in parthereinafter and in part will be obvious herefrom, or may be learned bypractice with the invention, the same being realized and attained bymeans of the methods, steps, parts, combinations and improvementspointed out in the appended claims.

Several processes, combinations and structural features individually andcollectively describe aspects of the cool-- ing techniques according tothe invention, it being understood that the invention consists win thenovel parts, constructions, arrangements, combinations, processes andsteps herein shown and described. Briefly and generally these includethe storage of a metal coolant, initially in solid form, in the annulusbetween the nozzle liner and outside casing. When the rocket is ignitedthere is a transfer of heat from the exhaust gas stream to the metalcoolant by way of the nozzle liner which functions as a heat exchanger,the coolant being disposed in optimum heat transfer relation with aportion of the liner. Upon melting, the cool-ant is caused to flowthrough a plurality of controlled passageways in the nozzle liner toform a succession of thin films which flow downstream within the gasboundary layer adjacent to the line-r to provide transpiration andevaporative film cooling. As films origin-ating upstream evaporate, theyare replenished by those formed further downstream.

In the illustrated embodiment the coolant flow is controlled by factorsincluding stagnation pressure of the combustion gases, (in liquidpropelled engines flow is thus responsive to engine throttling). Thecoolant is guided at low velocity along the boundary layer of the nozzlethroat in paths which include not only axial, but tangential componentsas well; a rotary component of movement is thus induced which tends tosupplement the coanda effect in avoiding direct interaction of the filmsand the gas stream.

Low mass flow rate of the coolant is provided by the disposition ofcoolant passageways by their geometry, and by choice of coolant. Thelatter may comprise a mechanically inert and readily fabricated metalsuch as aluminum or a binary solution of metals such as aluminum andmagnesium; the coolant preferably has a high enthalpy including highheat of vaporization and these properties are effectively translatedinto substantial heat absorption 'by the described transpirative andevaporati-ve film techniques.

The high density of the coolant contributes to low volumetric flow rateand with the above mentioned factors and others (high boiling point,high specific heat, low melting point) encourages the coolant to flow asa thin film which may be actually thinner than the nozzle boundary layerand thus isolated to some extent from direct interaction with the gasstream. In addition the film, being constantly replenished, serves toprovide a measure of protection for the liner against chemical attack bythe more reactive gases in the exhaust stream.

In several processes according to the invention the coolant flow iscontrolled in such manner as to leave residual coolant, which freezesduring a temporary shut down, in a region of the heat exchanging linerwhere sufficient heat will be available to remelt this coolant onrestart. Thus, utilization of substantially all coolant, and cyclicoperation of the coolant system, are provided.

In the processes according to the invention, certain ablative techniquesmay also be employed in the entrant portion of the nozzle.

Among the structural characteristics of the invention is the use of aspecial graphite arrangement as the nozzle liner. Illustratedembodiments employ pyrolitic or high density molded graphite formed as astack of graphite rings or annular members which are loaded incompression in the radial direction by the liquid coolant anddimensioned to define the requisite nozzle contour. The liner isprovided with a semi-porous character :for transpiration cooling bymeans of spaces between the rings which define supply channels for thecoolant flow. The arrangement is accordingly well suited to exploitationof the cooling processes of the invention, and moreover, facilitates theachievement of desired steady-state temperature conditions within thewall and in the coolant.

The use of graphite permits higher wall temperatures with a resultantminimization of heat flux. Thus less coolant is required. In addition,the graphite provides the desirable characteristic whereby strengthincreases with temperature.

Refractory metals such as tungsten are difiicult to work at roomtemperature because of their brittleness and high elastic modulus. In analternative embodiment this problem is overcome by a liner comprising afilamentary or fibrous refractory metal such as tungsten which iscompressed into the desired porous shape, eg an annular ring. Theporosity of the liner material may be readily and carefully controlledand may have metal coolant embodied therein.

The concept of stacking annular members to form the liner results inremoval of radial restraint and acts to relieve stresses in the liner;this condition is supplemented by the disposition of the metal coolantas a hydraulic or liquid pressure backing for the. liner to accommodatethermal-induced expansion of the graphite rings.

The stack arrangement with its integral channels also oifers aconvenient way of metering coolant fiow so that the coolant required toabsorb the heat flux of each ring is properly supplied and so that ascoolant film is depleted by evaporation, a regulated replenishmentthereof is constantly provided by coolant flowing into the boundarylayer via the passageways bet-ween the succeeding downstream rings.

In certain applications of the invention, the graphite liner mayadvantageously be provided with a reinforcement structure.

In several embodiments of the invention the coolant takes the form of amultilayer coil of metal wire wound around the liner. Besides itsadaptability to simple low cost fabrication techniques, this arrangementoffers the significant advantage of providing readily attained andcontrolled porosity manifested in the voids between adjacent turns whichprovide the required clearance volume and required distribution ofclearance. As noted below, this coolant configuration provides reducedcontact resistance and better utilization of the coolant storage space.

Other features and aspects of the invention which will become apparentupon the description below of certain exemplary embodiments and theirillustration in the drawings of which:

FIGURE 1 is an elevation view on reduced scale and partly in sectionillustrating the general outline of a cooled nozzle for asolid-propellant engine.

FIGURE 1A is an elevational sectional view on larger scale taken alongthe lines 1A1A of FIGURE 1;

FIGURE 1B is an elevational section-a1 view on larger scale taken alongthe lines 1B-1B of FIGURE 1;

FIGURE 1C is a fragmentary elevational detail view in sectionillustrating a portion of the structure illustrated in FIGURES 1A, 1B;

FIGURE 1D is a fragmentary sectional detail view illustrating amodification in the coolant feed system of FIGURES 1A, 1B;

FIGURE 1D is a fragmentary sectional detail view il' lustrating amodification in the coolant feed system of FIGURES 1, 1A, 1B, 1C;

FIGURE 1B is a fragmentary perspective detail view illustrating aportion of the system of FIGURES 1, 1A, 1B, 1C;

FIGURE 1F illustrates a modification to the liner of FIGURES 13;

FIGURE 2 is an elevational and sectional view taken along the lines 2-2of FIGURES l-3;

FIGURE 3 is an elevational sectional View taken along the lines 33 ofFIGURES 1A, 1B;

FIGURES 4 and 5 are elevational views partly in section and partlyschematic illustrating alternate embodiments of the nozzle structureillustrated in the preceding figures;

FIGURE 6 is a temperature profile diagram helpful in understandingcertain aspects of the invention;

FIGURE 7 is an elevational and sectional view of a nozzle-thrust chamberassembly for a liquid propellant rocket.

FIGURES 8 and 8A are framgentary elevational views of an alternate linerarrangement;

FIGURE 8B is a schematic elevational view;

FIGURE 8C is a detail fragmentary view, and FIG URE 8D an elevationalview in section, illustrating the formation and composition of the linerelements of FIG- URES 8, 8A; and

FIGURE 9 is an elevational and sectional view of the nozzle-thrustchamber assembly of FIGURE 7 with certain modifications thereto.

In the embodiment of FIGURES 1 to 3, the nozzle structure 1 is shownattached to a casing 2 which illus tratively encloses a thrust chamber3. The nozzle structure 1 includes a generally tubular heat exchangingliner 4 having a variable inner diameter for defining the requisitenozzle contour; in the example, this is a convergingdiverging or DeLaval type nozzle having a converging section in the nozzle entrance 5which is bounded by liner section 4a, a diverging section in the exitregion 6 which is bounded by liner section 40, and a throat section 7,therebetween, bounded by liner section 411.

The casing 1 is illustratively formed of two sections including aninsulation cover 10 of suitable composition such as an asbestos phenolicand a cap 11 forming a housing which may be of steel. The housing mayinclude a circ'urnferentially grooved flange 12 for facilitatingconnection of the nozzle as by a lock ring to the case 2 of the rocketstructure.

In the entrance region 5, between liner section 4a and cover 10 there isprovided a spacer 13 of insulating material such as graphite phenoliccloth tape. Illustratively, spacer 13 includes a rim section 14 seatedin a mating annular groove 15 of cover 10.

The space generally defined by the nozzle cover It! and the liner 4includes an annular coolant storage area 16 for retaining a metalcoolant such as an aluminum alloy embodied as a multi-turn multilayercoil of wire 17. As noted hereinbefore this arrangement avoids excesscontact resistance, provides good utilization of the storage space andprovides means for effectively controlling the coolant clearance volumeand the distribution thereof. Morover, by coating the wire or by usingadditional coils integrated in the spaces between coils, further controlover coolant behavior may be efiected.

For example, the potential disadvantage connected with the relativelylow vapor pressure of aluminum may be over-come by forming wire 17 of analuminum alloy which includes another metal, 61g. magnesium, forelevating the net vapor pressure of the resultant binary solution.Alternatively, the coil 17 may comprise aluminum'with an additional coilof magnesium disposed in the interstices thereof.

Extending through the inlet section 4a of liner 4 and through spacer 13is a passageway 14 which communicates with coolant chamber 16.Combustion pressure is thus applied to the chamber for controllingcoolant flow as described more fully hereinafter.

The annular liner sections 4a, 40 are formed of high density graphitesuch as ZTA or CARBONE P-5890. Section 4a, where preservation ofgeometry is not crucial, is preferably designed to ablate and itslength, together with the gases generated by ablation, augments thethickness of the exhaust boundary layer. The high heat of sublimation ofthe graphite matetrial results in relatively small losses in wallthickness. For relatively large nozzles, sections 4:: and 4c may beprovided with short duration ablative coating materials which ablateduring the initial coolant melt period.

The section 4b of liner 4 comprises a stack of annular graphite membersor rings, 4d, 42 preferably formed of pyrolytic or high density graphiteand having mating offset portions 19 for registration and stability.Rings 4d and 4e are disposed in alternate interlocked succession, 42,4d, 4e, 4d, etc., and may have different outside diameters so as topresent greater contact surface area and therefore better heat transferto coolant 17. The outer peripheries of rings 4d and dc, together withsection 40 may be provided with a layer of thin metal foil 18 alsohaving this effect when it melts.

The inner peripheral face of each member 4d is conveniently beveled toprovide a surface 4d, sloping downstream (see also FIG. 1e), while thecorresponding inner peripheral face of each member 4e, includes anintegral lip 4e on one edge thereof which is spaced from the adjacentbeveled turface 4d By this arrangement circumferential slots or channels20 sloping downstream are formed between each annular pair 4e, M.

In the upstream face of each member 4d an annular channel 4d isprovided, the open face of which is closed off by the abuttingdownstream face of the adjacent ring 4e. As may be clearly seen inFIGURE 1A, the resultant series of annular passageways or ducts 21 areinterconnected to form a coolant outlet header by registered bores 22 ineach ring. In addition, the upstream face of each ring 40. is providedwith a series of channels or grooves 411 (see also FIG. Ie), the openfaces of which are closed off by the abutting downstream face of theadjacent ring 4e. The channels 40% are routed angularly of the radial toprovide a tangential component therein and serve to interconnect thecoolant outlet header 21, 22 with the circumferentially slotted outlets20. Hence coolant supplied to header 21, 22 assumes the form of astacked series of interconnected annular sheets, each sheet feeding inturn the respective set of skewed channels 4:1 each set of channels 4:1jointly feeds in turn the respective annular outlet 20.

The inclination of the slots 20 guides the coolant in the downstreamdirection while the tangentially inclined channels 403 superimpose awhirling or vertical motion thereon. Channels 40' and outlets 20 aredimensioned for proper metering of coolant flow and in addition may becoated alone or together with other surfaces, with a suitable materialsuch as a carbide or oxide to inhibit clogging. Similar means may beadopted, or the coolant composition adjusted, to promote wetting.

As may be seen clearly in FIGURE 1A, the leading upstream ring 4dincludes a radial bore 25 which communicates with coolant feed pipes 26,27, preferably of graphite and disposed in the storage space 16. Pipe 27extends downstream and communicates with a pocket 16a in space 16 at itsaft, lowermost point via an inlet port 27a. Hence, the coolant space 16feeds the liner header 21, 22 from the continuously submerged pointwhere depleted coolant will collect under the action of thrust chamberpressure applied via passageway 14 (FIG- URE 1B) and the action ofgravity and acceleration forces in expected rocket attitudes. As shownin FIGURE 1d, feed pipe 27 may include a thermally controlled valveembodied as a plug 27b having a higher melting point than that of thecoolant whereby coolant flow is delayed until it reaches a temperatureabove its melting point.

In certain cases, the liner 4 may require structural reinforcement suchas shown in FIGURE 4 where the liner is backed by a refractory metalreinforcement 31.

The described nozzle structures may be fabricated as follows: In anexemplary process, the liner 4 is fabricated by assembling thepre-machined rings 4d, 4e, the fore and aft liner sections 4a, 40 andthe pre-molded spacer 13 in which the bore 14 is provided. Coolant 17,preceded where required by foil 18 or member 31, FIGURE 4, may then bewoundv under tension around the liner with the aid of a mandrel andappropriate jig. Lines 26 and 27 may be applied prior to the windingoperation or thereafter following the formation of bores therefor, inthe wound coolant. Thereafter insulation cover 10 is applied, preferablyby winding a tape form thereof around the coolant coil; the resultantassembly is fitted with cap 11.

Reference may be had to FIGURES 1-3 and FIGURE 6 in the followingdiscussion of exemplary operating conditions.

As suggested in the temperature profile of FIGURE 6, the operatingtemperature Tg of the combustion gas is assumed to be 8000 F. and thethrust chamber gas pressure, Pg 1000 p.s.i.a. It is also assumed thatthe specific heat ratio of the gas is about 1.16, the characteristicvelocity is 6600 ft./sec., and the molecular weight of the combustionproducts is about 21.

Following ignition at a time t a first transient interval occurs,lasting until a time 1 which in the example is about 4.1 seconds. At tthe temperature Tw, g at the liner wallgas interface is calculated to be4600 F. while the temperature Tw, c at the liner wall coolant interface,is 1250 F. This temperature is slightly above the melting point of thecoolant and the coolant starts to melt; the temperature Tw, i at thecoolant-insulation interface is F.

The second and third intervals may be regarded as steady state since theliner wall temperatures remain substantially constant until the engineis shut down or the coolant substantially depleted. Hence the heat fluxq will remain substantially constant during these intervals. During thesecond interval (e.g. 5.1 seconds), the temperature at thecoolant-insulation interface will rise to 1250 F. due to the absorptionby the coolant of the heat of fusion. During this second interval (t tot and during the first interval (t to t the coolant system is providingheat sink cooling.

At time t the third interval commences; all the coolant is liquifiedand, under the effects of external forces and the pressurizationprovided by the combustion gas via port 14, FIG. 1b, the coolantcommences to flow through lines 26, 27 to the liner header 21, 22, thechannels 4d and out of the slots 20. As noted hereinbefore the coolantis fed to the liner in a manner which provides a generally vorticalflow. The coolant film in an exemplary case is less than 0.001 inch andis controlled to be less than the gas boundary layer. Because of therelatively short distance between adjacent slots 20, the problem of filmstability is lessened and as previously noted, the rotary component ofmotion and the coanda effect, together with relatively low axialvelocity, results in the film being urged towards the liner wall. Theseeffects may be supplemented by suitable wetting agents Where required inthe form of a coolant ingredient or a liner coating. At the end of thethird interval (t all of the coolant is consumed.

In traversing the distance between adjacent slots 20, the coolant filmis evaporated but the liner is replenished with film at the succeedingslot where the coolant flow rate is metered to absorb the local heattransfer between channels. Further contributions to cooling by fluxreduction may result from a lowering of the heat transfer coeflicient atthe gas-liner interface due to the presence of the liquid film andbecause of the presence of the coolant vapors.

The liquid metal will reach the steady state temperature Tw, g, of thegas-liner interface as it approaches slots 20 and flows along the linerwall. It is preferred that the coolant under these conditions (e.g. T w,g:4600 F.) have a vapor pressure higher than the local gas pressure soas to insure complete vaporization and full endothermic utilization ofthe heat of vaporization. Thus, where aluminum is used as the coolant,it may be deployed as the solvent in a binary solution with a liquidmetal solute of higher vapor pressure such as magnesium, the solutionhaving the desired vapor pressure according to Raoults Law.

In the illustrated example, and based on the geometry of a typicalrelatively small nozzle, a coolant supply of about 6.8 lbs. will lastfor a period of about 96 seconds. At lower gas pressure, the coolantduration will increase, rising in the example to a period of about 250seconds for a chamber pressure of 300 p.s.i.a. Decreases in gastemperature, with wall temperatures unchanged, also extend the coolantsupply interval. At gas pressures of 1000 p.s.i.a. and 300 p.s.i.a., adecrease in gas temperature to 5500 F. will increase the coolant supplyduration to 300 seconds and 7 50 seconds, respectively.

The described operating conditions may include ablation of the linersection 4a and in some cases 40. Under the assumed conditions, withlocal flux conditions estimated to be 5 B.t.u./in. sec. and withablating section comprising a graphite material having a specific weightof 0.063 lb./ir1. and a heat of sublimation of 25,700 B.t.u./ 1b., anablation of one inch will occur in about 320 seconds.

During the described operating conditions, the liner 4 is subject tovarious stresses including pressure stresses resulting from theapplication of thrust chamber stagnation pressure to the coolant side ofthe liner. This pressure is larger than the internal gas pressure whichreduces as the local Mach number increases along the nozzle axis. Hencethe illustrated graphite liner is under compression and this conditionis favorable in the case of graphite which is stronger in compressionthan in tension and which becomes stronger as temperature rises.

The graphite members 4d, 4e are free to expand radially since they arebacked by the displaceable coolant. Hence thermal compressive stresseson the exhaust side of the liner which are associated with restraint arereduced. Thermal tensile stresses on the coolant side increase. However,pyrolytic graphic has good tensile properties as well and thus the lackof restraint makes better utilization of the material in exploiting itstensile properties while reducing the requirements for highercompressive strength.

In the exemplary operating mode, the liner wall on the coolant sideoperates at Tw,c=1250 F. If the temperature Tw,c of the wall can bemaintained at a higher value while the exhaust side, Tw g, remains thesame, thermal stresses will decrease. This may be accomplished by use ofthinner liner wall sections. The reduced wall thickness and temperaturegradient then cause thermal stresses to decrease to a greater extentthan pressure stresses increase. Thus net stress decreases. Furthermore,the increased temperature increases the graphite strength.

To raise the temperature Tw,c, the coolant flow may be 8 delayed beyondthe point Where its melting point is reached as by use of the plug 27b,FIG. 1d, which for some applications may be steel. Alternatively, thecoolant side of the liner may be provided with a thermal barrier, forexample, a coating of suitable material 30 such as zirconium oxide asshown in FIGURE 1F. In an exemplary design, the coating is about 0.002inch thick and has temperatures on the liner and coolant sides of 2700F. and 1250 F., respectively.

The techniques according to the invention are especially adapted forapplication to larger nozzles. For such larger nozzles, analysis willshow that the requisite coolant volurne occupies a portion only of theavailable annular space. It can also be shown that of the availableliner length, only a fraction thereof, e.g., to /2 throat radius, isneeded to provide the required heat exchange to melt the coolant. Thiscondition may be supplemented by providing the convergent and divergentliner sections with a short duration ablative coating operable duringthe coolant melt down period.

Although in larger nozzles the required coolant volume is relativelylessened as is the nozzle length necessary for heat exchange to melt thecoolant, a graphite liner in these cases (e.g. where throat diameter isgreater than about 16 inches and chamber pressure is about 1000p.s.i.a.) may require structural reinforcement since graphite propertiessuch as thermal conductivity, stress, allowable exhaust sidetemperature, and gas film heat transfer coefficient act to limit thegraphite liner wall thickness. If the wall is too thick the temperatureon its exhaust side will continue to rise above acceptable limits beforethe coolant melts and flow begins.

Schematically illustrating the foregoing and the weight reductionthereby achieved is the nozzle of FIGURE 5 which with the exceptionsnoted is intended to embody the features hereinbefore described. Thecoolant chamber 40 occupies a relatively small portion of the totalavailable annular space and is confined to the throat section 4b of theliner 4 where the requisite heat exchange is provided. The liner in thisembodiment serves primarily as a thermal barrier.

The chamber 40 comprises an annular space formed in the refractoryreinforcing structure 41 which also comprises the nozzle casing orcover. In the region of annular members 4d, 4e, the shroud member 41directly backs the liner but is protected from excess temperature by thecoolant flowing between members 4d, 4e. The remainer of the linerincluding divergent section 40 and convergent section 4a may comprise agraphite phenolic cloth backed by appropriate insulating material 43followed by reinforcement 41. In the throat section shroud 41 is incompression and may thus include fins 44 which serve as stiifeners andact also as heat transfer elements.

With the coolant chamber confined to the throat region, the expansioncone section 4c is freed of the higher compressive stress which wouldotherwise occur in this area. This section may also be cooled byradiation in high expansion ratio nozzles to provide further weightreduction, the decrease in internal gas pressure in these casespermitting a thinner shroud.

The embodiment of FIGURE 7 illustrates integration of the coolingtechniques in the design of a thrust cham ber for a liquid propellantrocket. The structure includes a single integrated flange 70facilitating connection of the chamber to the injector of a liquidpropellant non-regeneratively cooled rocket. With exceptions hereinafternoted, the cooling operations and assembly techniques are similar tothose described hereinbefore.

The embodiment of FIGURE 7 includes a liner 71 having a thrust chambersection 712?, convergent section 71a, throat section 7112 and divergingexit section 710. The leading sections 71t, 71a are designed to .ablateand are backed by an insulating layer 72 which may be of graphitephenolic cloth having vent holes (not shown) and which, through its lowthermal conductivity, insures 9 a low temperature at the outerinsulation jacket 73. The latter may comprise an asbestos phenolicmaterial molded around liner 71 and backing 72.

Serving as a housing as well as the means for retaining flange 70 is thecover 75. This cover is preferably of fibre glass filament woundconstruction and provides good structural backing for the linermaterials. Instead of serving additionally to secure flange 70, thelatter may be discarded and the fibre glass continued over a suitableportion of the injector.

The annulus between liner sections 71a, 71b, 71c and insulation jacket73, comprises the coolant storage area 76 for storing coolant 77 whichmay be as in the previously described embodiments. Additionally,structural reinforcement 78 may be provided.

The chamber 76 is pressurized by thrust chamber pressure via apassageway 79 bored through the liner 71 at the upstream edge ofconvergent section 71a, and through the contiguous sections of backing72 and jacket 73. In addition to eliminating the :pump requirements ofregeneratively cooled systems, this arrangement insures proper operationin vacuum conditions.

Liner section 71b comprises a stack of annular or ring members 714, 71earranged in an alternating pattern 71d, 71c, 71d, 71e, etc., asdescribed hereinbefore. The liner may include an integral header andfeed such as 21, 22, 26, 27 in previously described embodiments. In agravity free, vacuum condition the absence of drag insures constantacceleration during engine operation and the liquid coolant surface willlie generally normal to the thrust axis and recede downstream. In agravitational field at low thrust levels the liquid surface more nearlyparallels the thrust axis.

In an exemplary application of the system of FIGURE 7 where gastemperature is 8000 F. and thrust chamber pressure 300 p.s.i.a., acooling period of about 360 seconds may be realized under full thrustconditions of 3750 pounds, with an aluminum-magnesium coolant of about33 pounds. At 1000 p.s.i.a., 24 pounds of coolant would be required fora 360 second duration while 33 pounds of coolant would increase thecooling period to 480 seconds. Since the liquid metal coolant isultimately discharged with the jet stream it contributes to thrust, andthus only a fraction thereof need be regarded as inert.

The system of FIGURE 7 also functions effectively at reduced throttle.As throttle is reduced chamber pressure decreases; the required coolantflow rate decreases more rapidly than the actual flow rate so thatadequate coolant during reduced throttle is insured.

Duration of coolant supply increases for reduced throttle conditions.For example, operation at full throttle extends the coolant duration to19 minutes and if the coolant system were designed to this specificvalue, the coolant duration would be 38 minutes. The above per formancemay be improved by replacing the wire wound metal coolant with a castmetal configuration.

The system of FIGURE 7 offers major advantages over conventionalrefractory metal, regeneratively-cooled arrangements. The weight,density, and costly and diflicult working of refractory metals and theircomplicated double-wall structures .are eliminated, as is theconventional pump plumbing. Also eliminated is the need for the largerpump systems which would be required to satisfy cooling functions athigh temperatures such as 8000" F. In addition, the increasingavailability of coolant space with increased nozzle size renders thesystems according to the invention particularly adapted to larger sizerockets.

Stop-restart capability in liquid metal cooled systems presents seriousand unique problems. As the coolant is depleted, the residual supply isurged in any of several directions depending on the flight program.Following an engine shut down the residual metal coolant freezes;crystals are formed adjacent the lowest temperature area and growprogressively from this surface. The resultant solid mass may not be inproper :heat exchanging relationship with the liner either because it isnot in contact with the liner and/or, although in contact, it has beenso depleted that suflicient surface area of coolant is not available forthe necessary heat exchange.

If these conditions exist overheating and burnout can occur in thenozzle. To overcome these difficulties and to provide reliablestop-restart capability under the conditions of gravity-free flight orvertical flight in a gravitational field, the system of FIGURE 9 may beemployed.

As seen therein, the coolant storage volume is divided into a pluralityof chambers or enclosures 76, 76v by means which illustratively comprisean annular divider 80 coaxial with the thrust axis. Chamber 76v confinesresidual coolant to the area 71b of liner 71 where heat exchange ismaximum and this enclosure is fed from the main supply chamber 76 as bya longitudinal tube 81 which is preferably of molybdenum. Tube 81 has anoutlet 81a which communicates with residual chamber 76v via an orifice80a in divider 80 and an inlet 81b located in the after section of mainchamber 76.

Prior to initial take off, both enclosures 76 and 76v are provided withcoolant. Since the control over coolant porosity provided :by windingthe same in wire form is not available after the coolant has been oncemelted, this arrangement may be replaced by a coolant which is precastduring fabrication or added in molten form through a suitable aperture.

Upon initial engine firing, the coolant 77v in chamber 76v first meltsfollowed by the main mass of coolant 77. Coolant flows through channels71 between the members 71d, 71c, from the enclosure 76v to the linerslots 90. This supply is constantly replenished and enclosure 76vmaintained with coolant from the main supply 7 via 81b, 81, 81a and 80a.

Under the assumed flight conditions the main coolant supply 77 willrecede toward the aft end of the nozzle with its surface generallynormal to the thrust line; the residual enclosure 76v will remainfilled.

Following engine shut down both supplies 77 and 77v will graduallyfreeze. Although liner 71 is at a relatively high temperature, coolingthereof is effected due to evaporation of the coolant which occurs whenlocal gas pressure in the nozzle falls to ambient levels. This effect,together with the insulating properties of the jacket 73 which inhibitsheat transfer, promotes the formation of crystals of coolant 77v alongthe liner 71. Contraction and shrinkage tend in the direction of theliner and the over-all ecect is to leave the frozen coolant insatisfactory heat transfer relation with the liner notwithstanding somevoids will occur in various regions of the coolant mass.

Following engine restart, the rise in nozzle temperature liquefies thecoolant progressively, starting with the portion directly contiguous thenozzle throat. Until the main supply 77 is melted, pressurization viaport 79 is ineffective. However, combustion pressure is applied tocoolant chamber 76v via the leading channels 71 When the supply 77 hasmelted, the cooling operation stabilizes.

FIGURES 8 through 8D illustrate techniques involving the formation andapplication of a porous refractory metal, such as tungsten, as the linermaterial. This material forms a nozzle liner 98 on a primary refractorybackmg 99. As in previously illustrated configurations, liquid metalcoolant is deployed in contact with backing 99.

In the illustrated embodiment the elements of the liner 98 each comprisea porous annular ring 100 having the edges of its outer perimeterbevelled to form a series of annular channels 101 betweenadjacentelements. Each element also has at least one longitudinal notch 102formed in its outer periphery whereby, with the elements stacked and thegrooves 102 aligned, a header channel 103 is formed. This channel linksthe annular manifolds 101 for coolant distribution to the elements. Acoolant feedback system including a coolant tubular inlet 105'communicates with the distribution system 101, 103, by way of an openingin the primary liner structure 99.

An exemplary method of forming the elements 100 is illustrated in partin FIGURES 8B, 8C and 8D. Initially, a refractory wire 109 such astungsten is woven or knitted to form an appropriate configuration suchas the generally cylindrical knitted stockinette 110 shown in FIG- URES8B and 8C. This structure may then be partially compressed as by rollingthe ends of the stockinette towards the middle to form the doubletoroidal structure 111 shown in FIGURE 8D. This structure may then becompacted or compressed in a suitable die to form the annular rings 10%.The press forming operation provides the required nozzle contours andmay also impress the necessary bevels and grooves 102.

The porosity of the liner elements permits coolant to flow through themass thereof from the coolant supply system to the exhaust side of theliner wall whereby transpiration and evaporative film cooling isimplemented.

Control over the porosity or density of the annular rings 100 iscontrollable according to the nature or pattern of the knitting orWeaving, the wire diameter and the degree of pressing. Where requiredthe liner elements may also he formed with additional coolant flowchannels and the radial thickness of each element adjusted in accordancewith the porosity required at the position the element occupies alongthe liner wall. The density may also be adjusted by the insertion ofsegments of refractory metal fibers in the interstices of the mesh.

In the foregoing arrangement it is preferred that metal coolant besupplied integrally with the liner elements. This may be accomplished bya number of methods such as weaving coolant wire 115, FIG. 8C, into therefractory mesh, by adding the coolant in powder form prior to thepressing operation, by submersion of the liner elements in a bath of thecoolant, or by wrapping the coolant in foil form around the rolledstockinette prior to the compacting or pressing operation.

The illustrated and described techniques have been shown by way ofexample. Modifications Will undoubtedly occur to those skilled in theart. The invention is accordingly not limited to the specific techniquesshown and described but departures may be made therefrom Within thescope of the accompanying claims without departing from the principlesof the invention and without sacrificing its chief advantages.

What is claimed is:

1. An integrally cooled rocket engine nozzle capable of operation in thepresence of a rocket exhaust stream having temperatures substantiallyexceeding the melting points of refractory metals comprising a fusiblecoolant configured to permit expansion and having a vapor pressurecharacteristic for facilitating liquid film formation at rocketoperating temperatures, a nozzle liner bounding said exhaust stream anddimensioned to function as a heat exchanger for forming a fused film ofcoolant on the exhaust stream side of said liner, said liner including astack of annular rings, a casing enclosing said liner with a storagespace therebetween for storing said fuzible coolant, said storage spacebeing oriented relative to said liner in such a manner as to convey heatof fusion to said coolant via said liner, and passageways in said stackinterconnecting said storage space and the exhaust stream side of saidliner and dimensioned and located for continuously transporting saidfused coolant to said exhaust stream side of said liner to provideevaporative film cooling thereof.

2. A nozzle as defined in claim 1 including a pressurizing passagewaybetween the exhaust stream side of said liner and said storage space forsupplying exhaust stream pressure to said coolant.

3. A nozzle as defined in claim 1 in which said rings comprise agraphite material and have spaces therebe tween for defining at least aportion of said passageways.

4. A nozzle as defined in claim 1 in which said rings comprise a porousfilamentary refractory metal and a metal coolant.

5. A nozzle as defined in claim 1 in which said storage space includes amain storage area and residual coolant storage space, the latter beingin heat exchanging relationship with said liner and meansinterconnecting said residual coolant storage space with said mainstorage area.

6. A liner as defined in claim 1 in which said rings are adapted to beloaded in compression and comprise pyrolitic graphite.

7. A liner as defined in claim 1 in which said rings include a thermalbarrier coating on their exteriors.

8. A liner as defined in claim it in which said rings comprisefilamentary refractory wire compacted to a predetermined porosity andhaving a coolant material integrated therein.

9. A liner as defined in claim 8 in which said rings include a poroustungsten structure, said porous tungsten structure being impregnatedwith a coolant comprising at least two different metals adapted whenliquified to form a binary solution for controlling the vapor pressureof said coolant.

10. An integrally cooled rocket engine nozzle capable of operation inthe presence of a rocket exhaust stream having temperaturessubstantially exceeding the melting points of refractory metalscomprising a fusible coolant, a nozzle liner bounding said exhauststream and configured to form a heat exchanger for continuously fusingsaid coolant, said liner including an annular graphite member, a casingenclosing at least a portion of said liner with a storage spacetherebetween for storing said coolant, pressure transmitting meanspressurizing said storage space from said exhaust stream, a fuziblecoolant manifold passageway within said storage space, and transpiration passageways in said liner communicating with said manifoldpassageway and the exhaust stream side of said liner, said coolantcomprising a composition having a vapor pressure characteristicfacilitating film forming for causing evaporative films to How alongsaid exhaust stream side of said liner.

11. A nozzle as defined in claim 10 including a pressurizing passagewaybetween the exhaust stream side of said liner and said storage space forsupplying exhaust stream pressure to said coolant whereby the flow ofsaid coolant is responsive to throttling of said engine.

12. A nozzle as defined in claim 10 in which said liner comprises astack of annular members in contact with said coolant and free to expandin the presence of thermal forces for providing stress relief therein.

13. A nozzle as defined in claim 10 in which said graphite memberincludes a stack of annular graphite members having channelstherebetween for forming said passageways, said channels includingregions oriented to impart an axial and rotary movement to said coolant.

14. A nozzle as defined in claim It in which said storage space includesa main storage area and a residual coolant storage space, the latterbeing in heat exchanging relationship with said liner, and meansinterconnecting said residual coolant storage space with a region insaid main storage area where depleted coolant will collect under varyingflight conditions.

15. A nozzle as defined in claim 10 in which said graphite member isbacked by a refractory reinforcing structure and in which saidreinforcing structure includes an integral displaced section formingsaid casing.

16. A nozzle as defined in claim 10 in which said graphite member isbacked by a metallic foil.

17. A nozzle as defined in claim 10 in which said coolant comprises acoil of wire.

18. A nozzle as defined in claim 10 including an annular divider in saidstorage space for dividing said space into a main and residualenclosure, said residual enclosure being in heat transfer relation withsaid liner and in communication with said main enclosure near the exitregion of said nozzle.

19. A nozzle as defined in claim 10 in which said coolant comprises acombination of at least two metals forming a binary solution when fused.

References Cited by the Examiner UNITED STATES PATENTS Ward 6035.6Kimmel 6035.6 Feldman 6039.66 X Vest 60-39.66 Warnken 6035.6 Hsia6039.66 Terner 6039.66 X Prosen 6039.66 X

CARLTON R. CROYLE, Primary Examiner.

1. AN INTEGRALLY COOLED ROCKET ENGINE NOZZLE CAPABLE OF OPERATION IN THE PRESENCE OF A ROCKET EXHAUST STREAM HAVING TEMPERATURES SUBSTANTIALLY EXCEEDING THE MELTING POINTS OF REFRACTORY METALS COMPRISING A FUSIBLE COOLANT CONFIGURED TO PERMIT EXPANSION AND HAVING A VAPOR PRESSURE CHARACTERISTIC FOR FACILITATING LIQUID FILM FORMATION AT ROCKET OPERATING TEMPERATURES, A NOZZLE LINER BOUNDING SAID EXHAUST STREAM AND DIMENSIONED TO FUNCTION AS A HEAT EXCHANGER FOR FORMING A FUSED FILM OF COOLANT ON THE EXHAUST STREAM SIDE OF SAID LINER, SAID LINER INCLUDING A STACK OF ANNULAR RINGS, A CASING ENCLOSING SAID LINER WITH A STORAGE SPACE THEREBETWEEN FOR STORING SAID FUZIBLE COOLANT, SAID STORAGE SPACE THEREBETWEEN FOR STORING SAID FUZIBLE LINER IN SUCH A MANNER AS TO CONVEY HEAT OF FUSION TO SAID COOLANT VIA SAID LINER, AND PASSAGEWAYS IN SAID STACK INTERCONNECTING SAID STORAGE SPACE AND THE EXHAUST STREAM SIDE OF SAID LINER AND DIMENSIONED AND LOCATED FOR CONTINUOUSLY TRANSPORTING SAID FUSED COOLANT TO SAID EXHAUST STREAM SIDE OF SAID LINER TO PROVIDE EVAPORATIVE FILM COOLING THEREOF. 